The invention concerns a method for the manufacture of a fiber-reinforced thermoplastic composite module, for example, a shell element of an aircraft fuselage, and a device for the execution of the method.
Fiber-reinforced composite modules, such as shell elements of an aircraft fuselage, have in a manner of known art a backing structure formed from a multiplicity of longitudinal stiffeners, which are directly connected to a skin field, and a multiplicity of circumferential stiffeners, which are connected to the skin field via fittings, e.g., clips, and in addition are supported on the longitudinal stiffeners via supporting elements, e.g., cleats.
The manufacture of such fuselage segments is undertaken using either a differential form of construction or an integral form of construction. In the differential form of construction the individual components are produced separately from one another and are then assembled to form the complete module. What is particularly disadvantageous in this form of construction is the time- and cost-intensive assembly of the complete module. In addition the differential form of construction requires a multiplicity of connecting elements for purposes of connecting the module components together. The integral form of construction, in which the individual components are produced in an integral manner as a complete module, features a reduced assembly effort and, by virtue of the increase in the level of integration, a greater production effort. For purposes of reducing the production effort the patent U.S. Pat. No. 6,613,258 B1 proposes, for example, the manufacture of a fuselage barrel from fiber-reinforced thermoplastic composite materials. Here cured longitudinal stiffeners are laid in depressions of a cylindrical core. The core is then set in rotation about its longitudinal axis and a web-type laminate is wound onto the core. The laying down of the laminate is undertaken via a laying down head with the application of temperature and pressure, as a result of which the thermoplastic matrix is in a molten state and the laminate is welded securely to the longitudinal stiffeners. Disadvantageous in this solution, however, are the very tight production tolerances on the fuselage barrels.